1. Field of the Invention
The present invention generally relates to gas turbine engines and, more particularly, to rotor members for such gas turbine engines.
2. Description of the Prior Art
At present, most of the world's small aircraft gas turbine engines incorporate a multi-stage titanium compressor. Typically, the first stages are axial compressors and the last stage is a one-piece centrifugal impeller. Typical compressor delivery temperatures are around 850° F. (at the tip of the impeller) and the bore air temperature is around 600° F. The resulting delta T between the two locations is roughly 250° F. With this delta T, the industry has been able to achieve adequate low cycle fatigue (LCF) lives.
Recent demands to improve the engine fuel consumption have resulted in the compressor delivery temperature increasing to 1000° F. and greater. Unfortunately, this high compressor delivery temperature creates high temperature regions on the rear cavity of the impeller. Titanium, at these high temperatures, suffers from creep/fatigue interaction as well as oxidation problems which lowers the LCF life to an unacceptable value.
Accordingly, it has been proposed to replace the titanium impeller with a nickel impeller that does not suffer from the creep/fatigue and oxidation issues at these temperatures. However, the physical properties of nickel, with the presence of the delta T in excess of 400° F., results in an LCF which is again commercially not viable.